A method of manufacturing a cooled gas turbine component includes forming a core with an outer surface. The outer surface includes a core feature. The method also includes casting an outer wall of an airfoil about the core. The outer wall has an exterior surface and an interior surface. The interior surface includes a shaped inlet portion that corresponds to the core feature. Moreover, the method includes forming an outlet portion through the outer wall to fluidly connect the outlet portion to the shaped inlet portion. The shaped inlet portion and the outlet portion cooperatively define a cooling aperture through the outer wall.
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11. A cooled gas turbine component for a gas turbine engine comprising:
an airfoil;
an outer wall of the airfoil, the outer wall having an exterior surface and an interior surface, the outer wall having a thickness measured between the exterior surface and the interior surface, the outer wall including a constant-thickness portion at which the thickness is constant; and
a cooling aperture that extends along an axis through the constant-thickness portion of the outer wall, the cooling aperture including:
a shaped inlet portion included on and recessed relative to the interior surface; and
an outlet portion having an outlet axis that is straight and linear, extending through the outer wall, and fluidly connected to the inlet portion at an intersection, the outlet portion having a diameter that remains constant from the exterior surface to the intersection;
wherein the inlet portion continuously surrounds the outlet portion about the axis to widen the cooling aperture about an entire periphery thereof at the interior surface, the inlet portion having an inlet width and a depth, wherein the inlet width of the inlet portion tapers and reduces along the depth of the inlet portion and fluidly connects to the outlet portion at the intersection; and
wherein the inlet width of the inlet portion at the intersection is greater than the diameter at the intersection to define a step at the intersection.
1. A cooled gas turbine component for a gas turbine engine comprising:
an airfoil;
a cast outer wall of the airfoil, the outer wall having an exterior surface and an interior surface, the outer wall having a thickness measured between the exterior surface and the interior surface, the outer wall including a constant-thickness portion at which the thickness is constant; and
a cooling aperture that extends along an axis through the constant-thickness portion of the outer wall, the cooling aperture having a length measured along the axis, the cooling aperture being a partly-cast aperture along the length and being a partially-machined aperture along the length, the cooling aperture including:
a cast inlet portion included on and recessed relative to the interior surface; and
a machined outlet portion that is straight and linear along the axis, that is fluidly connected to the inlet portion at an intersection, and that extends through the outer wall to the exterior surface, the outlet portion having a diameter that remains constant from the exterior surface to the intersection;
wherein the inlet portion has a width and a depth, wherein the width of the inlet portion gradually reduces along the depth of the inlet portion toward the outlet portion; and
wherein the inlet width of the inlet portion at the intersection is greater than the diameter at the intersection to define a step at the intersection.
2. The cooled gas turbine component of
wherein the inlet portion is at least partially conic.
3. The cooled gas turbine component of
4. The cooled gas turbine component of
5. The cooled gas turbine component of
6. The cooled gas turbine component of
7. The cooled gas turbine component of
8. The cooled gas turbine component of
9. The cooled gas turbine component of
10. The cooled gas turbine component of
further comprising an internal channel that extends through the internal wall and that fluidly connects the first cavity and the second cavity; and
wherein a first width of the cooling aperture is smaller than a second width of the internal channel.
13. The cooled gas turbine component of
wherein the outlet axis extends at an acute angle relative to the exterior surface of the outer wall.
14. The cooled gas turbine component of
15. The cooled gas turbine component of
16. The cooled gas turbine component of
17. The cooled gas turbine component of
wherein the outlet portion has an outlet surface with a machined surface finish.
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This application is a divisional of U.S. patent application Ser. No. 15/266,481, filed Sep. 15, 2016, the entire disclosure of which is incorporated by reference herein.
The present disclosure generally relates to a gas turbine component and, more particularly, relates to a gas turbine component with a cooling aperture having a shaped inlet and a method of forming the same.
Gas turbine engines are generally used in a wide range of applications, such as aircraft engines and auxiliary power units. In a gas turbine engine, air is compressed in a compressor, mixed with fuel, and ignited in a combustor to generate hot combustion gases, which flow downstream into a turbine section. In a typical configuration, the turbine section includes airfoils, such as stator vanes and rotor blades, disposed in an alternating sequence along the axial length of a generally annular hot gas flow path. The rotor blades are mounted at the periphery of one or more rotor disks that are coupled in turn to a main engine shaft. Hot combustion gases are delivered from the engine combustor to the annular hot gas flow path, thus resulting in rotary driving of the rotor disks to provide an engine output.
Due to the high temperatures in many gas turbine engine applications, it is desirable to regulate the operating temperature of certain engine components, particularly those within the mainstream hot gas flow path in order to prevent overheating and potential mechanical issues attributable thereto. As such, it is desirable to cool the airfoils of the rotor blades and stator vanes to prevent or reduce oxidation, thermo-mechanical fatigue, and/or other adverse impacts to the airfoil. Mechanisms for cooling turbine airfoils include ducting cooling air through internal passages and then venting the cooling air through holes formed in the airfoil. However, given the high temperature of engine operation, cooling remains a challenge.
Accordingly, it is desirable to provide gas turbine engines with improved airfoil cooling. Furthermore, other desirable features and characteristics of the present invention will become apparent from the subsequent detailed description of the invention and the appended claims, taken in conjunction with the accompanying drawings and this background of the invention.
In one embodiment, a method of manufacturing a gas turbine component for a gas turbine engine is disclosed. The method includes forming a core with an outer surface. The outer surface includes a core feature. The method also includes casting an outer wall of an airfoil about the core. The outer wall has an exterior surface and an interior surface. The interior surface includes a shaped inlet portion that corresponds to the core feature. Moreover, the method includes forming an outlet portion through the outer wall to fluidly connect the outlet portion to the shaped inlet portion. The shaped inlet portion and the outlet portion cooperatively define a cooling aperture through the outer wall.
In another embodiment, a cooled gas turbine component for a gas turbine engine is disclosed. The gas turbine component includes an airfoil. Also, the gas turbine component includes an outer wall of the airfoil. The outer wall has an exterior surface and an interior surface. Also, the gas turbine component includes a cooling aperture that extends through the outer wall. The cooling aperture includes a cast inlet portion included on the interior surface. The cooling aperture also includes an outlet portion extending through the outer wall and fluidly connected to the inlet portion. The inlet portion has a width and a depth, and the width of the inlet portion gradually reduces along the depth of the inlet portion toward the outlet portion.
The present disclosure will hereinafter be described in conjunction with the following drawing figures, wherein like numerals denote like elements, and wherein:
The following detailed description is merely exemplary in nature and is not intended to limit the present disclosure or the application and uses of the present disclosure. Furthermore, there is no intention to be bound by any theory presented in the preceding background or the following detailed description.
Broadly, exemplary embodiments disclosed herein include gas turbine engines with turbine components having improved cooling characteristics. Methods of manufacturing the turbine components are also disclosed. In particular, exemplary embodiments include turbine airfoils with an outer wall having at least one cooling aperture with a shaped inlet. The shaped inlet may increase flow rate through the cooling aperture for improved cooling. Additionally, methods of manufacturing the turbine component may include casting the shaped inlet and subsequently forming an outlet of the cooling aperture through the casting to fluidly connect the outlet to the shaped inlet. Additionally, in some embodiments, the cooling aperture may be adjusted (e.g., by widening the hole and/or changing the axis of the outlet) to change flow characteristics of the cooling aperture. Other details of the present disclosure will be discussed below.
The gas turbine engine 100 has an overall construction and operation that is generally understood by persons skilled in the art. The gas turbine engine 100 may be disposed in an engine case 101 and may include a fan section 120, a compressor section 130, a combustion section 140, a turbine section 150, and an exhaust section 160, which are arranged sequentially along a longitudinal axis 180. As used herein, the term “axial” refers to a direction generally parallel to the longitudinal axis 180. A radial axis 190 is also included in
The fan section 120 may include a fan, which draws in and accelerates air. A fraction of the accelerated air from the fan section 120 is directed through a bypass section 170 to provide a forward thrust. The remaining fraction of air exhausted from the fan is directed into the compressor section 130.
The compressor section 130 may include a series of compressors that raise the pressure of the air directed into it from the fan section 120. The compressors may direct the compressed air into the combustion section 140.
In the combustion section 140, the high pressure air is mixed with fuel and combusted. The combusted air is then directed into the turbine section 150.
The turbine section 150 may include a series of rotor assemblies 192 and stator assemblies 194, both of which are represented schematically in
Within the turbine section 150, the rotor assemblies 192 may include a plurality of rotor blades 200, an example embodiment of which is illustrated in
To allow the turbine section 150 to operate at desirable elevated temperatures, certain components are cooled. For example, in some embodiments, the rotor blade 200 may include cooling apertures that include one or more features of the present disclosure. In additional embodiments, one or more of the stator assemblies 194 (e.g., an airfoil of the stator assembly 194) may include the cooling apertures of the present disclosure. As will be discussed, the cooling aperture may include a shaped inlet portion. The shaped inlet portion provides increased flow for improved cooling. Manufacturing techniques are also discussed below for providing the shaped inlet portion to one or more of these turbine components.
The rotor blade 200 includes an airfoil 202, a platform 204, and a root 206. The platform 204 is configured to radially contain turbine airflow within a shroud of the turbine section 150. The root 206 extends from the underside of the platform 204 and is configured to couple the blade 200 to a turbine rotor disc (not shown). In general, the rotor blade 200 may be made from any suitable material, including high heat and high stress resistant aerospace alloys, such as nickel based alloys, Mar-M-247, single crystal materials, directionally solidified materials, or others.
The airfoil 202 projects radially outwardly from the platform 204 and terminates at a blade tip 220. The airfoil 202 is formed by a body 208 with an outer wall 209. The outer wall 209 may include a first portion 210 and a second portion 212 that cooperate to define an airfoil shape. The first portion 210 of the outer wall 209 defines a pressure side with a generally concave shape, and the second portion 212 of the outer wall 209 defines a suction side with a generally convex shape. In a chordwise direction, the portions 210, 212 of the outer wall 209 are joined at a leading edge 214 and trailing edge 216. As used herein, the term “chordwise” refers to a generally longitudinal dimension along the airfoil 202 from the leading edge 214 to the trailing edge 216.
As noted above, the rotor blade 200, particularly the airfoil 202, may be subject to extremely high temperatures resulting from high velocity hot gases ducted from the combustion section 140 (
The cooling apertures 222 may extend through the outer wall 209. It will be appreciated that the number of cooling apertures 222 and the arrangement of the cooling apertures 222 may vary without departing from the scope of the present disclosure. The cooling apertures 222 may be relatively small and arranged generally in rows and/or columns on the airfoil 202, proximate the leading edge 214, the blade tip 220, and/or other areas of the airfoil 202. As shown in
In some embodiments, the cooling apertures 222 may provide film cooling to the blade 200. Specifically, the cooling apertures 222 may be arranged to provide a cooling film of fluid onto the exterior surface (i.e., the hot side surface) of the airfoil 202.
As discussed above, the airfoil 202 may include the outer wall 209 with first and second portions 210, 212 joined at the leading edge 214 and the trailing edge 216. As shown in
Also, as noted above, the airfoil 202 may include the cooling system 221, which cools the airfoil 202. Various features of the cooling system 221 will be described in detail below according to exemplary embodiments.
It will be appreciated that, although the cooling system 221 is shown and described in relation to the airfoil 202 of the rotor blade 200, one or more features of the cooling system 221 may be incorporated within another area of the rotor blade 200 and/or within another turbine component. For example, the cooling system 221 may be incorporated within a stator assembly 194 of the turbine section 150 of the engine 100 without departing from the scope of the present disclosure.
In exemplary embodiments, the cooling system 221 may form part of a high efficiency, multi-walled turbine airfoil cooling arrangement or a serpentine airfoil cooling arrangement. Other cooling systems 221 may be provided, including those of different wall and cavity structures.
The cooling system 221 may include one or more internal voids 234 defined within the airfoil 202. The internal void 234 may comprise a cavity, a passageway, a channel, a pocket, a hollow, or other void within the airfoil 202. As shown in the embodiment of
The cooling system 221 may additionally include the cooling apertures 222 mentioned above. One of the cooling apertures 222 is included in
The cooling aperture 222 may be in fluid communication with the internal void 234 in some embodiments. For example, the inner end 248 may be fluidly connected and open to the forward cavity 236 in some embodiments. Thus, cooling fluid may flow from the forward cavity 236, into the inner end 248 of the cooling aperture, and out of the airfoil 202 via the outer end 246. Thus, the inner end 248 may define an inlet of the cooling aperture 222, and the outer end 246 may define an outlet of the cooling aperture 222.
As shown in
The cooling aperture 222 may have a length 254 that is measured along (e.g., parallel to) the axis 252. The length 254 may be measured from the outer end 246 to the inner end 248 of the cooling aperture 222.
The cooling aperture 222 may also have a width 256 (e.g., diameter). The width 256 may be measured transverse to (e.g., perpendicular to) the axis 252 between opposing areas of an inner surface 250 of the aperture 222.
There may be several differences between the cooling apertures 222 and other features of the cooling system 221. Specifically, the cooling apertures 222 may extend through the exterior surface 228 of the airfoil 202, whereas the forward cavity 236, rear cavity 238, and internal channel 240 are enclosed within the airfoil 202. Also, the width 256 of the cooling apertures 222 may be substantially smaller than the width 244 of the internal channel 240. Specifically, whereas the internal channel 240 may have a width 244 of at least 0.040 inches, the cooling aperture 222 may have a width 256 between approximately 0.012 to 0.030 inches. Also, the cooling apertures 222 may provide cooling (e.g., film cooling) to the exterior surface 228, whereas the forward cavity 236, rear cavity 38, and/or internal channel 240 may provide cooling to the interior surface 230.
As shown in
As shown in
The inlet portion 262 may be defined by a concave inlet surface 270. The inlet surface 270 may extend along the axis 252 between an inner rim 269 and the intersection 266. In some embodiments, the width 256 of the inlet portion 262 may vary along its length from the inner rim 269 to the intersection 266. For example, the width 256 may gradually reduce (e.g., taper) from the inner rim 269 to the intersection 266. Stated differently, the width 256 of the inlet portion 262 at the intersection 266 may be substantially equal to the width 256 of the outlet portion 264 at the intersection as shown in
The inlet surface 270 may be defined according to a variety of shapes. For example, the inlet surface 270 may be at least partially conic (e.g., frustoconic) in some embodiments. As such, the inlet surface 270 may have two-dimensional curvature. In other embodiments, the inlet portion 262 may be concave and generally dome-shaped, hemispherical, or otherwise rounded. As such, the inlet surface 270 may have three-dimensional curvature.
In some embodiments, the frustoconic inlet portion 262 may be substantially centered about the axis 252. Thus, as shown in
In some embodiments, the outlet portion 264 may comprise the majority of the length 254 of the cooling aperture 222. Also, the length of the outlet portion 264 (measured from the outer rim 267 to the intersection 266 in
The inlet portion 262 may be open to the forward cavity 236 within the airfoil 202. The outlet portion 264 may be in fluid communication with the inlet portion 262. Therefore, a fluid flowpath may be defined from the forward cavity 236, to the inlet portion 262, and out of the airfoil 202 via the outlet portion 264 for film cooling of the airfoil 202.
The inlet portion 262 may define a widened and shaped inlet portion of the cooling aperture 222. The inlet portion 262 may provide a gradual transition along the interior surface 230 to the outlet portion 264. In some embodiments, the inlet portion 262 may define a chamfer or chamfer-like feature for the outlet portion 264 of the cooling aperture 222. Thus, the profile of the inlet surface 270, the tapering width 256 of the inlet portion 262, and/or other features may increase fluid flow through the cooling aperture 222. As a result, the airfoil 202 may be cooled more efficiently and effectively. Also, because the inlet portion 262 increases fluid flow, the width 256 of the outlet portion 264 may be reduced, making the outer wall 209 stronger and more robust.
Methods of manufacturing the airfoil 202 will now be discussed with reference to
As shown in
As shown in
Next, the method 300 of
It will be appreciated that the core 414 may be formed in other ways as well. For example, additive manufacturing techniques (e.g., 3D printing) may be employed for forming the core 414. These techniques may also be used to form the core feature 416 of the core 414.
Subsequently, the method 300 may continue at 306, in which a tool 418 is fabricated. As shown in
Next, at 308, the core 414 may be disposed within the cavity 424 as shown in
Subsequently, at 310 of the method 300, a shell material 429, such as wax, may be provided within the region 428. The shell material 429 may be hardened to define a shell 430 about the core 414. Accordingly, a first intermediate article 432 may be formed that includes the core 414 and the shell 430.
The method 300 may continue at 311. At 311, the first intermediate article 432 may be removed from the cavity 424 of the tool 418, and the first intermediate article 432 may be dipped one or more times in a slurry material 434. The slurry material 434 may be a ceramic material. As shown in
Then, at 312 of the method 300, the airfoil 202 may be cast about the core 414 and within the region 442. Specifically, as shown in
Then, as shown in
Subsequently, at 316 of the method 300, the outlet portion 264 of the cooling aperture 222 may be formed. For example, as shown in
During the development process of a cooled vane or blade, the orientation and diameter of the cooling aperture 222 may change based on test results. There may also be tolerances on where the intersection 266 is located on the inner surface 230 (i.e., where the outlet portion 264 breaks out on the inner surface 230). Accordingly, in some embodiments (for example, in which the airfoil 202 is to be tested before in-flight use), the method 300 may continue at 318. At 318, the airfoil 202 may be tested, for example, to determine the flow and cooling characteristics of the cooling aperture 222. In some embodiments, the airfoil 202 may be tested in connection with CFD modeling techniques and tools.
Next, the method 300 may continue to 320, wherein it determined whether to adjust the cooling aperture 222 to provide more desirable flow and cooling characteristics. If no adjustments are needed, the method 300 may finish. If, on the other hand, the testing of 318 indicates that adjustments are needed for the cooling aperture 222, then the method may continue to 322.
At 322, it may be determined whether it is necessary to re-cast the airfoil 202. In some cases, such as the embodiment of
In other embodiments of 322 of the method 300, it may be necessary to re-cast the airfoil 202 (i.e., 322 answered positively). Thus, the method 300 may continue to 326, wherein the core 414 is re-formed. The core 414 may be re-formed using the same core die 400 formed at 302 and illustrated in
It will be appreciated that the adjustments shown in
Referring now to
As shown in
It will be appreciated that the manufacturing method 300 discussed above may be used to manufacture the thickened area 470″ and other areas of the outer wall 209″. For example, at 312 of the method 300, the outer wall 209″ may be cast about a core 414″, and the inner surface 230″ (including the thickened area 470″) may be formed according to the outer surface 415″ of the core 414″. The core 414″ may include a recess 472″, and the core feature 416″ may project outwardly from the recess 472″. Thus, the thickened area 470″ may be formed inversely according to the surfaces of the recess 472″. Likewise, the inlet portion 262″ may be formed inversely according to the surfaces of the core feature 416″.
Accordingly, the shaped inlet portions 262 of the cooling apertures 222 of the airfoil 202 provide several advantages. The shaped inlet portion 262 improves the flow coefficient of the cooling aperture 222, allowing more air to pass through a given hole width 256. For an aperture 222 that is perpendicular to the wall 209, the shaped inlet portion 262 can increase flow at least 15%, to 40%. Moreover, in cases in which the shaped inlet portion 262 extends continuously about the axis 252 of the outlet portion 264, the flow can be less sensitive to variations in the direction of flow and wall thickness variations. Additionally, the cooling apertures 222 can be less likely to plug with dirt or other debris because the inlet portion 262 is tapered instead of having a sharp inlet edge. This is because flow through the cooling apertures 222 is less likely to result in recirculation zones, which can cause particle build-up. Moreover, the shaped inlet portions 262 can direct and channel the air, minimizing flow separations within the aperture. Accordingly, film cooling can occur in an effective manner.
Furthermore, manufacturing of the airfoil 202 can be completed in an efficient manner. The casting and subsequent drilling or EDM methods described above can be completed in a controlled fashion for high manufacturing accuracy. The methods provide time savings as well. Also, testing and adjusting the cooling apertures 222 may be completed in a convenient manner because, instead of having to form another core with new tooling, molds, etc., the outlet portion 264 can be adjusted by re-drilling as discussed above.
While at least one exemplary embodiment has been presented in the foregoing detailed description, it should be appreciated that a vast number of variations exist. It should also be appreciated that the exemplary embodiment or exemplary embodiments are only examples, and are not intended to limit the scope, applicability, or configuration of the present disclosure in any way. Rather, the foregoing detailed description will provide those skilled in the art with a convenient road map for implementing an exemplary embodiment of the present disclosure. It is understood that various changes may be made in the function and arrangement of elements described in an exemplary embodiment without departing from the scope of the present disclosure as set forth in the appended claims.
Patent | Priority | Assignee | Title |
Patent | Priority | Assignee | Title |
10605095, | May 11 2016 | General Electric Company | Ceramic matrix composite airfoil cooling |
4565490, | Oct 26 1978 | ALSTOM SWITZERLAND LTD | Integrated gas/steam nozzle |
4992025, | Oct 12 1988 | Rolls-Royce plc | Film cooled components |
5059093, | Jun 07 1990 | United Technologies Corporation | Compressor bleed port |
5222617, | Oct 17 1990 | Rolls-Royce plc | Drilling turbine blades |
5700131, | Aug 24 1988 | United Technologies Corporation | Cooled blades for a gas turbine engine |
6241468, | Oct 06 1998 | Rolls-Royce plc | Coolant passages for gas turbine components |
6616405, | Jan 09 2001 | MITSUBISHI HITACHI POWER SYSTEMS, LTD | Cooling structure for a gas turbine |
7059825, | May 27 2004 | RTX CORPORATION | Cooled rotor blade |
7237595, | Oct 29 2003 | Siemens Aktiengesellschaft | Casting mold |
8522557, | Dec 21 2006 | Siemens Aktiengesellschaft | Cooling channel for cooling a hot gas guiding component |
8628292, | Mar 28 2007 | SIEMENS ENERGY GLOBAL GMBH & CO KG | Eccentric chamfer at inlet of branches in a flow channel |
8858176, | Dec 13 2011 | FLORIDA TURBINE TECHNOLOGIES, INC | Turbine airfoil with leading edge cooling |
9057277, | Dec 06 2011 | RTX CORPORATION | Systems, devices, and/or methods for producing holes |
9249679, | Mar 15 2011 | GE INFRASTRUCTURE TECHNOLOGY LLC | Impingement sleeve and methods for designing and forming impingement sleeve |
9296039, | Apr 24 2012 | RTX CORPORATION | Gas turbine engine airfoil impingement cooling |
9394798, | Apr 02 2013 | Honeywell International Inc. | Gas turbine engines with turbine airfoil cooling |
20090285683, | |||
20110132562, | |||
20110162387, | |||
20130156602, | |||
20140050938, | |||
20140219778, | |||
20160061451, | |||
20170306764, | |||
20180274370, | |||
EP992654, | |||
EP2777842, | |||
EP2995773, | |||
WO2013037662, | |||
WO2016022140, |
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